Virtual Shock Shaping Using Microjet Arrays

Abstract

Arrays of high momentum microjets are used to generate single or multiple oblique shocks in the supersonic cross-flow. The shocks generated using microjets can be tailored to be parallel or coalescing depending upon the application. Flow visualization using Shadowgraph and density field measurements using Background Oriented Schlieren technique were obtained for a range of test conditions. The results obtained using the two methods are consistent and show a linear variation of oblique shock angle with microjet supply pressure up to the range tested. The density field obtained using BOS clearly shows the oblique shock generated using microjet, the jump in density across the shock, the extent of high density field and the expansion fan. There is a large historical interest towards high speed commercial travel and military combat . The high speed combat aircraft have been quite successful but one of the biggest challenges in the establishmen t of commercial supersonic travel involves dealing with the impact of associated shock system, well known as Sonic boom. Many technologies in the past have been, and are being developed to reduce the signature of sonic boom. DARPA’s Shaped Sonic Boom Demonstration (SSBD) program 1-2 on F-5E aircraft demonstrated the reduction in sonic boom signature by extensively modified forward fuselage and the effort was followed up by NASA’s Shaped Sonic Boom Experiment (SSBE) program. Gulfstream’s Quite Spike program 3-4 utilized a long telescopic fuselage extension to alter the N-wave pressure signature on F-15 aircraft. The design of a supersonic inlet has been and will continue to pose another technological challenge due to its complex flow field with multiple shocks. Inlet ramps, single or multi are offering a good percentage of pressure recovery up to certain velocities but there performance degrade at higher Mach number operations. Isentropic ramps are complicated to design and expensive to manufacture and appear sporadically in aircraft history. In traditional inlets in supersonic aircraft, the flow is decelerated through a series of oblique shock waves, which eventually terminate in a normal shock followed by a subsonic diffuser. The interaction of these shocks, e specially the terminal normal shock, results in Shock Wave Boundary Layer Interactions (SWBLI) inside the inlet. Such SWBLIs generally lead to degradation in the inlet performance in terms of: significant pressure losses; flow unsteadiness in the duct and on the compressor fan- face; local -or in severe cases, global separation and non-uniform flow. Consequently, the control of SWBLI for supersonic inlets is highly desirable for the present and next generation of aircraft. Traditionally this has been achieved through the use of (active or passive) bleed systems on the inlet surfaces 5-6 . Although this approach has been shown to reduce some of the SWBLI related adverse effects, the cost of such systems in terms additional weight, drag and complexity makes